CN108228921B - Wind tunnel test device and design method thereof - Google Patents
Wind tunnel test device and design method thereof Download PDFInfo
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Abstract
The embodiment of the invention provides a wind tunnel test device and a design method thereof, relates to a wind tunnel test examination technology, and can solve the problem that a plurality of state examinations are difficult to perform in a single wind tunnel test. The test assessment device of the wind tunnel has the characteristic of multistage compression, and a line segment L of a surface molded line of a multistage compression area bent from an N segment1、L2、L3……、LnConstitution L0The flow direction length of a test section with a zero attack angle is obtained, the stepped heat flow distribution characteristic stable in the flow direction is obtained in the same train number wind tunnel test through the idea of continuously and gradually changing the attack angle, the response of the material in various thermal states can be checked in the train number test, the pureness of the flow field is considered, the multistage compression oblique shock waves are intersected at one point, and the interference of the multi-channel shock wave reflection on the flow field is avoided. The defects of low efficiency, long period and high cost of the original material examination test are overcome, the test period is greatly shortened, and the engineering application value is high.
Description
Technical Field
The invention relates to the technical field of wind tunnel test examination, in particular to a wind tunnel test device and a design method thereof.
Background
The wind tunnel test can provide a simulated environment close to a real flight condition, and is one of effective means for examining the structure/thermal protection material. A large number of wind tunnel tests need to be carried out on the structure and thermal protection of the hypersonic aircraft, and especially in the demonstration of a scheme and the screening stage of materials, the ablation resistance, the oxidation resistance and the thermal insulation performance of the materials need to be thoroughly examined. Generally speaking, the wind tunnel test has long preliminary preparation time and low efficiency, and the test of one train number can only check one state, thus bringing huge period cost and expenditure cost to the development process, forming the bottleneck of scheme demonstration and restricting the progress of the whole research process.
Disclosure of Invention
The invention aims to overcome the defects of the prior art, provides a wind tunnel test device and a design method thereof, and can solve the problem that a plurality of state assessments are difficult to perform in a single wind tunnel test.
The technical solution of the invention is as follows:
on the one hand, the wind tunnel test piece examination device has the characteristic of multi-stage compression, and a line segment L formed by bending N sections of surface molded lines of multi-stage compression regions1、L2、L3……、LnConstitution L0The flow direction length L of each stage of compression area is calculated according to the required wall surface heat flow by using the principle that the flow direction length of a test section with a zero attack angle and the intersection points of oblique shock waves generated by a plurality of stages of compression areas are the same and adopting a planar oblique shock wave front-rear flow parameter relation0、L1、L2、L3……、LnAnd angle of attack alpha corresponding to each stage of compression region1、α2、α3……、αn。
In another aspect, a method for designing the above apparatus includes the following steps:
firstly, preliminarily selecting a spray pipe according to test assessment conditions, and calculating the air flow parameters at the outlet of the spray pipe by using a one-dimensional constant adiabatic flow parameter basic relational expression according to the geometrical parameters of the spray pipe and the incoming flow conditions, wherein the geometrical parameters comprise outlet area, outlet-throat area ratio, the incoming flow conditions comprise total temperature and total pressure, the test assessment conditions comprise enthalpy and heat flow, and the air flow parameters at the outlet comprise pressure, temperature and Mach number;
secondly, calculating the length and the attack angle of a compression area according to a relation formula of flow parameters before and after the plane oblique shock wave and on the premise of meeting test examination conditions and taking the intersection points of the oblique shock wave of the nuclear surface to be examined as the same by taking a required wall examination heat flow value as a target and obtaining the length L of the compression area1、L2、L3……、LnAngle of attack alpha1、α2、α3……、αn；
Thirdly, establishing a wind tunnel three-dimensional solid model containing a multi-stage compression region according to the selected geometrical size of the spray pipe, carrying out grid division, setting boundary conditions and initial conditions, carrying out simulation by adopting computational fluid dynamics software to obtain the wall surface heat flow distribution and the wind tunnel internal flow field distribution of a test section, adjusting the length and the attack angle of each stage of compression region according to a numerical simulation result, and meeting the heat flow specified by test assessment conditions, wherein the geometrical size of the spray pipe comprises a throat area, an expansion angle and an outlet area;
fourthly, debugging and verifying a wind tunnel test;
and fifthly, finally determining the length of each stage of compression area and the geometric parameters of the attack angle, and determining the scheme of the assessment device.
The principle of the multi-state wind tunnel test assessment device and the design method thereof provided by the embodiment of the invention is that a steady stepped heat flow distribution characteristic along the flow direction is obtained on the surface of a test piece in the same train number wind tunnel test through the concept of continuously and gradually changing the attack angle, the response of materials in various thermal states can be assessed in the train number test, the purity of a flow field is considered, multi-stage compression oblique shock waves are crossed at one point, and the interference of multi-channel shock wave reflection on the flow field is avoided. The defects of low efficiency, long period and high cost of the original material examination test are overcome, the test period is greatly shortened, and the engineering application value is high.
Drawings
The accompanying drawings, which are included to provide a further understanding of the embodiments of the invention and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the principles of the invention. It is obvious that the drawings in the following description are only some embodiments of the invention, and that for a person skilled in the art, other drawings can be derived from them without inventive effort.
FIG. 1 is a schematic view of a wind tunnel test device with a test section including a compression area;
FIG. 2 is a schematic view of a multi-stage compression test apparatus;
FIG. 3 is a flow chart of the design of a multi-stage compression test apparatus.
X in the figure0、X1、X2、X3Is the compression point flow direction position; l is0The length of the flow direction of the test section with the attack angle of zero; l is1、L2、L3Respectively the flow direction lengths of the compressed areas of the test sections; alpha is alpha1、α3、α3Respectively is an attack angle of a test section compression area; in addition, only 3 compression regions are shown in fig. 2, and in practical application, the number of compression stages can be adjusted as required.
Detailed Description
Specific embodiments of the present invention will be described in detail below with reference to the accompanying drawings. In the following description, for purposes of explanation and not limitation, specific details are set forth in order to provide a thorough understanding of the present invention. However, it will be apparent to one skilled in the art that the present invention may be practiced in other embodiments that depart from these specific details.
It should be noted that, in order to avoid obscuring the present invention with unnecessary details, only the device structures or processing steps closely related to the scheme according to the present invention are shown in the drawings, and other details not so relevant to the present invention are omitted.
In order to enable the wind tunnel test of a single train number to examine various test states, the efficiency is greatly improved, and the period and the cost are reduced. The embodiment of the invention provides a wind tunnel test device, as shown in figure 1, the wind tunnel test device comprises a spray pipe 1, a test section 2, a multi-stage compression region 3, a diffusion section 4 and a vacuum region 5, the wind tunnel test device has the characteristic of multistage compression, the molded line of the outer surface of the multistage compression region consists of N sections of bent line segments L0, L1, L2, L3 … … and Ln, L0 is the flow direction length of a test section with zero attack angle, based on the principle that the intersection points of oblique shock waves generated in a plurality of sections of compression regions are the same, a relation formula of front and back flow parameters of the plane oblique shock waves is adopted, the flow length L0, L1, L2, L3 … …, Ln of each stage of compression region and the corresponding attack angle α 1, α 2, α 3 … …, α n of each stage of compression region are calculated according to the required wall heat flow, and the relationship of the flow parameters before and after the plane oblique shock wave is a calculation method well known to those skilled in the art, and will not be described herein.
The principle of the multi-state wind tunnel test assessment device and the design method thereof provided by the embodiment of the invention is that a steady stepped heat flow distribution characteristic along the flow direction is obtained on the surface of a test piece in the same train number wind tunnel test through the concept of continuously and gradually changing the attack angle, the response of materials in various thermal states can be assessed in the train number test, the purity of a flow field is considered, multi-stage compression oblique shock waves are crossed at one point, and the interference of multi-channel shock wave reflection on the flow field is avoided. The defects of low efficiency, long period and high cost of the original material examination test are overcome, the test period is greatly shortened, and the engineering application value is high.
A design method of the wind tunnel test device comprises the following steps:
firstly, preliminarily selecting a spray pipe according to test examination conditions, and calculating the air flow parameters at the outlet of the spray pipe by using a one-dimensional constant adiabatic flow parameter basic relational expression, wherein the parameters comprise pressure, temperature, Mach number and the like, and the one-dimensional constant adiabatic flow parameter basic relational expression is a calculation method well known by the technical personnel in the field and is not described more herein;
secondly, designing the length and the attack angle of a compression area according to a relation formula of flow parameters before and after the plane oblique shock wave by taking a given wall surface heat flow value as a target and taking the intersection points of all oblique shock waves as a principle on the premise of meeting the test requirement to obtain the length (L) of the compression area1、L3、L3) Angle of attack (alpha)1、α3、α3) Geometric parameters are equal; the selection of parameters is according to the examination heat flow examination requirements (Q) of different compression surfacesw) Establishing, nozzle outlet parameter pressure (p)1) Density (p)1) Temperature (T)1) Mach number (M)1) Several oblique shock wave intersection points are selected to be far away from the surface with the check core as far as possible for known quantity, so that the flow field is ensured to be free of interference. The specific parameter solution is carried out according to the following formula:
the flat plate heat flow density calculation formula is as follows:
wherein
1) St is the number of the Stantons,
s is a comparison factor, CFIs a flat wall friction resistance coefficient
T*The total temperature of the wall surface is,
T*＝Te+0.5(Tw-Te)+0.22(Tr-T∞) (4)
Cf is notThe laminar flow n1 is-0.12, n2 is 0.5, and C1 is 1.328, which is the friction resistance coefficient of a flat plate wall surface where the incompressible gas flows; for turbulent flow n 1-0.648, n 2-0.2, C1-0.074
Reynolds number with plate length as characteristic scale:
calculated as laminar flow, S-0.80, calculated as turbulent flow, S-0.82
2)qwFor the heat flow density, the values are given according to the experimental requirements,
3)ρe、verespectively are the density and the speed of the outer airflow of the boundary layer behind the oblique shock wave,
Meis the Mach number, R, of the oblique shock wave flowgIs a gas constant, TeIs the gas temperature after the oblique shock wave. Rhoe、Me、TeCalculated according to the oblique shock wave relational expression
4)cpFor constant specific heat at constant pressure, given in the design according to the physical property constant
5) Tw is the wall temperature and is set to 300K
6) Tr is the recovery temperature of the liquid crystal,
r is a rewarming factor, 0.85 is taken for laminar flow and 0.89 is taken for turbulent flow
It can be seen that in the calculation formula of the heat flux density, besides the wall surface heat flux, other parameters need to be according to the gas parameter M after the oblique shock wavee、ρe、TeCalculating with wall temperature, knowing the parameter p of incoming flow1、ρ1、T1、M1
Mach number after laser
Beta and delta are respectively a shock wave angle and an air flow break angle, wherein the air flow break angle is the same as the break angle of the flat plate
Density ratio of front to back of oblique shock wave
Temperature ratio before and after oblique shock wave:
the relation between the laser angle and the air flow break angle is as follows:
by formulas (8) to (11), M can be expressede、ρe、TeExpressed as a single-valued function of the shock angle beta, and expressed by equations (2) to (7), equation (1) is expressed as containing only the heat flux density qwWall temperature TwA shock angle beta, a plate length L, a constant specific heat Cp,And (3) an expression of the dynamic viscosity coefficient mu and the wave front airflow parameters, wherein the expression only contains an unknown shock wave angle beta, and the shock wave angle beta is obtained by solving the expression. Then, according to the formula (11), the air flow break angle delta can be solved, and further the break angle alpha of the flat plate can be obtained.
Thirdly, establishing a wind tunnel three-dimensional solid model containing a compression region according to the geometric structure of the wind tunnel test device, carrying out grid division, setting boundary conditions and initial conditions, carrying out simulation by adopting computational fluid dynamics software to obtain the wall surface heat flow distribution and the wind tunnel internal flow field distribution of the test section, and adjusting the geometric parameters such as the length of the compression region, the attack angle and the like according to the numerical simulation result;
fourthly, carrying out debugging and verification of a wind tunnel test;
and fifthly, finally determining the length of each stage of compression area and the geometric parameters of the attack angle, and determining the scheme of the assessment device.
The wind tunnel test device provided by the embodiment relates to a process of checking surface heat flow of a part, obtains stable stepped heat flow distribution characteristics along a flow direction in the same train number wind tunnel test through a thought of continuously and gradually changing an attack angle, can check response of materials in various thermal states in the train number test, considers the purity of a flow field, crosses multi-stage compression oblique shock waves at one point, and avoids interference of multi-channel shock wave reflection on the flow field. The defects of low efficiency, long period and high cost of the original material examination test are overcome, the test period is greatly shortened, and the engineering application value is high.
Features that are described and/or illustrated above with respect to one embodiment may be used in the same way or in a similar way in one or more other embodiments and/or in combination with or instead of the features of the other embodiments.
It should be emphasized that the term "comprises/comprising" when used herein, is taken to specify the presence of stated features, integers, steps or components but does not preclude the presence or addition of one or more other features, integers, steps, components or groups thereof.
The many features and advantages of these embodiments are apparent from the detailed specification, and thus, it is intended by the appended claims to cover all such features and advantages of these embodiments which fall within the true spirit and scope thereof. Further, since numerous modifications and changes will readily occur to those skilled in the art, it is not desired to limit the embodiments of the invention to the exact construction and operation illustrated and described, and accordingly, all suitable modifications and equivalents may be resorted to, falling within the scope thereof.
The invention has not been described in detail and is in part known to those of skill in the art.
Claims (2)
1. The wind tunnel test device is characterized by comprising a spray pipe, a test section, a multi-stage compression region, a diffusion section and a vacuum region, and the wind tunnel test device has the multi-stage compression characteristic that the surface molded line of the multi-stage compression region consists of N +1 segment line sections L0、L1、L2、L3……、LnThe method comprises the steps of calculating the flow direction length L of each stage of compression region according to the required wall surface heat flow by using the principle that the intersection points of oblique shock waves generated by the multi-stage compression regions are the same and adopting a planar oblique shock wave front-rear flow parameter relation0、L1、L2、L3……、LnAnd L1、L2、L3……、LnAngle of attack alpha corresponding to each stage of compression region1、α2、α3……、αnWherein L is0Is the test segment flow direction length with zero angle of attack.
2. A method for designing a wind tunnel test device according to claim 1, comprising the steps of:
firstly, preliminarily selecting a spray pipe according to test assessment conditions, and calculating spray pipe outlet airflow parameters by using a one-dimensional constant adiabatic flow parameter basic relational expression according to spray pipe geometric parameters and incoming flow conditions, wherein the spray pipe geometric parameters comprise an outlet area, an outlet-throat area ratio, the incoming flow conditions comprise total temperature and total pressure, the test assessment conditions comprise an enthalpy value and heat flow, and the outlet airflow parameters comprise pressure, temperature and Mach number;
secondly, calculating the length and the attack angle of a compression area according to a relation formula of flow parameters before and after the plane oblique shock wave and on the premise of meeting test examination conditions and taking the intersection points of the oblique shock wave of the nuclear surface to be examined as the same by taking a required wall examination heat flow value as a target and obtaining the length L of the compression area0、L1、L2、L3……、LnAngle of attack alpha1、α2、α3……、αn；
Thirdly, establishing a wind tunnel three-dimensional solid model containing a multi-stage compression region according to the selected geometrical size of the spray pipe, carrying out grid division, setting boundary conditions and initial conditions, carrying out simulation by adopting computational fluid dynamics software to obtain the wall surface heat flow distribution and the wind tunnel internal flow field distribution of a test section, adjusting the length and the attack angle of each stage of compression region according to a numerical simulation result, and meeting the heat flow specified by test assessment conditions, wherein the geometrical size of the spray pipe comprises a throat area, an expansion angle and an outlet area;
fourthly, debugging and verifying a wind tunnel test;
and fifthly, finally determining the length of each stage of compression area and the geometric parameters of the attack angle, and determining the scheme of the assessment device.
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Citations (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102607799A (en) * | 2012-02-10 | 2012-07-25 | 南京航空航天大学 | Device for changing Mach number in supersonic velocity wind tunnel model experiment and working method |
| CN103366052A (en) * | 2013-06-27 | 2013-10-23 | 中国航天空气动力技术研究院 | Method for analyzing hot pneumostatic elasticity of hypersonic aerocraft |
| CN103592100A (en) * | 2013-10-30 | 2014-02-19 | 中国运载火箭技术研究院 | Grid wing wind tunnel test model scaling method |
| CN104461677A (en) * | 2014-10-30 | 2015-03-25 | 中国运载火箭技术研究院 | Virtual thermal test method based on CFD and FEM |
| CN104897406A (en) * | 2015-06-04 | 2015-09-09 | 大连海事大学 | Rotary stamping compressed rotor plane cascade experimental system |
| WO2016003765A1 (en) * | 2014-06-30 | 2016-01-07 | Airborne America, Inc. | Improved wind tunnel design with expanding corners |
| CN105468814A (en) * | 2015-10-29 | 2016-04-06 | 南京航空航天大学 | Helicopter flight characteristic computing method considering entire-helicopter aerodynamic interaction |
| CN106021831A (en) * | 2016-07-26 | 2016-10-12 | 厦门大学 | Design method of self-adaptive communication reverse reflux tank air inflow channel |
| CN106197936A (en) * | 2016-06-24 | 2016-12-07 | 中国航天空气动力技术研究院 | Aircraft head-shield two lobe rotating separation wind tunnel test methods and device |
| CN106202804A (en) * | 2016-07-22 | 2016-12-07 | 北京临近空间飞行器系统工程研究所 | Complex appearance aircraft distributed heat ambient parameter Forecasting Methodology based on data base |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP3757269B2 (en) * | 2001-08-24 | 2006-03-22 | 独立行政法人 宇宙航空研究開発機構 | Method and apparatus for reducing pressure fluctuation in wind path in recirculating supersonic wind tunnel |
-
2016
- 2016-12-13 CN CN201611144664.0A patent/CN108228921B/en active Active
Patent Citations (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102607799A (en) * | 2012-02-10 | 2012-07-25 | 南京航空航天大学 | Device for changing Mach number in supersonic velocity wind tunnel model experiment and working method |
| CN103366052A (en) * | 2013-06-27 | 2013-10-23 | 中国航天空气动力技术研究院 | Method for analyzing hot pneumostatic elasticity of hypersonic aerocraft |
| CN103592100A (en) * | 2013-10-30 | 2014-02-19 | 中国运载火箭技术研究院 | Grid wing wind tunnel test model scaling method |
| WO2016003765A1 (en) * | 2014-06-30 | 2016-01-07 | Airborne America, Inc. | Improved wind tunnel design with expanding corners |
| CN104461677A (en) * | 2014-10-30 | 2015-03-25 | 中国运载火箭技术研究院 | Virtual thermal test method based on CFD and FEM |
| CN104897406A (en) * | 2015-06-04 | 2015-09-09 | 大连海事大学 | Rotary stamping compressed rotor plane cascade experimental system |
| CN105468814A (en) * | 2015-10-29 | 2016-04-06 | 南京航空航天大学 | Helicopter flight characteristic computing method considering entire-helicopter aerodynamic interaction |
| CN106197936A (en) * | 2016-06-24 | 2016-12-07 | 中国航天空气动力技术研究院 | Aircraft head-shield two lobe rotating separation wind tunnel test methods and device |
| CN106202804A (en) * | 2016-07-22 | 2016-12-07 | 北京临近空间飞行器系统工程研究所 | Complex appearance aircraft distributed heat ambient parameter Forecasting Methodology based on data base |
| CN106021831A (en) * | 2016-07-26 | 2016-10-12 | 厦门大学 | Design method of self-adaptive communication reverse reflux tank air inflow channel |
Non-Patent Citations (5)
| Title |
|---|
| "2009年国外高超声速技术发展概述";张冬青 等;《飞航导弹》;20100915(第9期);第2-11页 * |
| "三级压缩椎导乘波体设计技术与实验分析";吕侦军 等;《实验流体力学》;20151015;第29卷(第5期);摘要第38页、第1章39-40页 * |
| "高超声速进气道发展研究";徐锦 等;《战术导弹技术》;20160826;第25-32页 * |
| "高超声速飞行器攻角和侧滑角测量算法";马小龙 等;《指挥控制与仿真》;20150815;第36-40页 * |
| Bruno Goffert et al."Numerical Study of Wall Ventilation in a Transonic Wind Tunnel".《Journal of Aerospace Technology and Management》.2015,第81-92页. * |
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